A Qualitative Piloted Evaluation of the Tupolev Tu-144 Supersonic Transport

NASA / TM-2000-209850

Robert A. Rivers and E. Bruce Jackson Langley Research Center, Hampton, Virginia

C. Gordon Fullerton and Timothy H. Cox Dryden Flight Research Center, Edwards, California

Norman H. Princen Boeing Commercial Airplane Group, Long Beach, California

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A Qualitative Piloted Evaluation of the Tupolev Tu-144 Supersonic Transport

Robert A. Rivers and E. Bruce Jackson Langley Research Center, Hampton, Virginia

C. Gordon Fullerton and Timothy H. Cox Dryden Flight Research Center, Edwards, California

Norman H. Princen Boeing Commercial Airplane Group, Long Beach, California

National Aeronautics and Space Administration

Langley Research Center Hampton, Virginia 23681-2199

February 2000 Acknowledgments The U.S. team would like to thank the entire Tupolev and IBP teams for helping us complete these flight tests both ahead of schedule and with a high degree of quality. We would like to extend a special thanks to the following Tupolev and IBP personnel for their exceptional contributions:

Prof. Alexander Pukhov _po_eccop AnexcaH_p _yXOB Edgar Krupyanski 3_rap KpynsHcnM_ Vladimir Sysoev Bna_MMHp C_COeB Sergei Borisov Cepre_ BopMcoB Victor Pedos BHKTOp _e_oc Anatoli Kriulin AHaTon_ Kp_ynMH Mikhail Melnitchenko MHxaMn MensHMqeH_o Yuri Tsibulin 10pM_ _m6ynMH

Available from:

NASA Center for AeroSpace Information (CASI) National Technical Information Service (NTIS) 7121 Standard Drive 5285 Port Royal Road Hanover, MD 21076-1320 Springfield, VA 22161-2171 (301) 621-0390 (703)605-6000 Table of Contents Nomenclature ...... 1 Introduction ...... 2 Aircraft Description ...... 2 Aircraft Geometry ...... 2 Aerodynamic Surfaces ...... 3 Propulsion System ...... 4 Fuel System ...... 5 Hydraulic System ...... 5 Electrical System ...... 6 Fire Detection and Extinguishing Systems ...... 7 Air Conditioning and Pressurization Systems ...... 7 Anti-Ice System ...... 8 Navigation and Communications ...... 8 Manual Flight Control System ...... 9 Autopilot System ...... 14 Landing Gear ...... 14 Cockpit Layout ...... 15 Crew Arrangement ...... 15 Flight Equipment ...... 15 Flight Test Planning & Preparation ...... 17 Method of Tests ...... 17 Planning Process ...... 17 Flight Readiness Review and Flight Task Examination ...... 18 Pilot Briefing ...... 18 Flight Monitoring and Control ...... 18 Flight Test Summaries ...... 20 Flight 21 - September 15, 1998 ...... 20 Flight 22 - September 18, 1998 ...... 21 Flight 23 - September 24, 1998 ...... 22 Observed Vehicle Characteristics ...... 23 Ground Handling ...... 23 Thrust Management ...... 23 Takeoff/Cleanup ...... 23 Handling Qualities ...... 24 Acceleration/Climb ...... 24 Supersonic Cruise ...... 25 Descent ...... 25 Structural Characteristics ...... 25 Low Speed Characteristics ...... 26 Traffic Pattern ...... 26 Landing/Ground Roll ...... 27 Conclusions ...... 27 References ...... 29 Appendix A: Description of Maneuvers ...... 30 Appendix B. Extended narratives for each U.S. piloted flight ...... 32

Abstract

Two U.S. research pilots evaluated the Tupolev Tu-144 supersonic transport aircraft on three dedicated flights: one subsonic and two supersonic profiles. The flight profiles and maneuvers were developed jointly by Tupolev and U.S. engineers. The vehicle was found to have unique operational and flight characteristics that serve as lessons for designers of future supersonic transport aircraft. Vehicle subsystems and observed characteristics are described as are flight test planning and ground monitoring facilities. Maneuver descriptions and extended pilot narratives for each flight are included as appendices.

Nomenclature da asymmetric elevon (aileron) deflection MAC mean aerodynamic chord de symmetric elevon (elevator) deflection MCP Mode Control Panel

AC alternating current N1 engine fan speed

ACM air cycle machine N3 engine turbine speed

ADI Attitude Director Indicator PID parameter identification

AGL above ground level PIO pilot-induced oscillation

AOA angle of attack PLA power lever angle (throttle setting)

APU Auxiliary Power Unit S Laplace operator

A/T autothrottle SPI Sensitive Pitch Indicator

CG center of gravity RPM revolutions per minute DC direct current TACAN Tactical Air Navigation DME Distance Measuring Equipment T.E. trailing edge EGT exhaust gas temperature UHF Ultra High Frequency FE flight engineer

g 1 decision speed HSR High Speed Research

V 2 safety speed ICS Interphone Communication System VHF Very High Frequency IDG Integrated Drive Generator

glof liftoff speed ILS Instrument Landing System VOR Very High Frequency Omnidirectional Range IMN indicated Mach number v rotation speed INS Inertial Navigation System VRI Vertical Regime Indicator ITB Integrated Test Block

LG landing gear

1 Introduction tim quantitative test results to tim qualitative pilot evaluations, handling qualities criteria boundaries Under the auspices of tim NASA High-Speed can be set for design of future supersonic transports. Research (HSR) program, an out-of-service Tu-144 supersonic transport aircraft was re-engined, refur- . Carry out qualitative handling qualities evaluations bished and proven as an experimental supersonic fly- of the Tu-144 using multiple pilots. The maneuvers ing laboratory by Tupolev Aircraft Company (Tupolev for tl_e Tu-144 evaluation represented normal op- ANTK). Nineteen flights of the modified aircraft erational maneuvers for this type of aircraft. The U.S. (Tu-144LL) were flown by Tupolev research pilots in team, witll the help of Tupolev, defined tim maneu- 1996-1997 and data for six flight experiments were ac- vers for tiffs testing. Where possible each evaluation quired. pilot flew and evaluated tl_e same set of maneuvers at the same flight conditions in order to account for A subsequent eight-flight program, flown in 1998, pilot variation in tl_e evaluation process. included tllree flights during which an evaluation of the handling qualities of tim Tu-144LL was performed This report represents a summary of tim aircraft by NASA research pilots. This report describes the con- systems, handling qualities, and flight characteristics duct and results of tllese three flights (known as flights of tim Tu-144LL aircraft. Also covered are the experi- 21, 22 and 23). ment and flight preparation processes that were used in conducting these tl_ree test flights. Participants in the flight program included IBP Aircraft, Ltd. (as exclusive contractual representative The information on Tu-144 systems was obtained of Tupolev), Tupolev ANTK, and tim U.S. Piloted over a two-week period of instruction by Tupolev en- Evaluation Team: representatives from NASA Dryden gineers and pilots in a lecture format. No written docu- Flight Research Center (DFRC), NASA Langley Re- ments were available to tim U.S. team with the excep- search Center (LaRC), and tile Boeing Company. The tion of tim untranslated Russian flight manual. The sys- flight experiments were performed jointly at Tupolev's tem lectures were in Russian and were interpreted by flight research facility near Zhukovsky, Russia. translators provided to tile U.S. team. Considering tiffs, the information contained below may have minor dis- The objectives of tile follow-on eight flights were: crepancies, but it can be considered accurate to the degree of providing a general description of system op- 1. Provide tim United States witl_ first-hand piloting eration. experience of the Tu- 144LL Supersonic Flying Labo- ratory. The Tu-144LL was developed to support Aircraft Description flight research beneficial to tile development of a next generation supersonic transport aircraft. The Tu-144LL Serial Number 77114 (figure 1) is a Tu-144LL is one of only two supersonic transport modified Tu-144D aircraft with newer engines neces- aircraft types tllat are flying in tile world today. To sitating nacelle modifications (fully described in tim maximize the benefit of the Tu- 144LL to U.S. avia- following section). Aircraft 77114 was built in 1981 tion, members of tile piloting community needed to and was tile final aircraft off tile production line. It has understand the characteristics and capabilities of tl_e only been used as a research aircraft, accumulating research airplane. The pilots' understanding of tile approximately 83 hours prior to being placed in non- airplane will greatly aid tile planning of future aero- flyable storage in 1990. In 1993 tile aircraft was brought nautical experiments using tile Tu-144LL, and per- out of storage and modifications were commenced lead- haps will lead to experiments relating to operational ing to the "LL" or "Flying Laboratory" designation and requirements for future high-speed civil transport a return to flight on November 29, 1996. aircraft. Aircraft Geometry 2. Collect additional supersonic quantitative handling qualities data for tl_e Tu-144 aircraft. By comparing The aircraft is a delta planfonn, low wing, four- engine supersonic transport, witla a retractable canard

2 Figure1.Tu-144LLFlyingLaboratory andhingednose,asshownin figure2. Thenominal fuselage. The canard includes both leading- and trail- cockpitcrewconsistsof twopilots,anavigatorseated ing-edge flaps that deflect when the canard is deployed; betweenthepilots,andaflightengineer.Generalspeci- they are faired when the canard retracts. These devices ficationsof theaircraftarelistedin Table1. are not actuated as part of the control system, but are deployed into a fixed position when the canard is ex- Aerodynamic Surfaces tended. The canard provides additional lift forward of the center of gravity to aid in reducing the takeoff and In addition to the main wing, a unique retractable landing approach pitch angle and reducing the trailing- canard is located just aft of the cockpit on top of the edge-up deflection of the elevons during low-speed 215'6"

19'10"

Figure 2. Three-view drawing of Tu-144

3 Table1.Tu-144LLVehicleCharacteristics

Length 65.7m (215ft 6in.) Wingspan 28.8m (94ft. 6 in.) Height,wheelsup 12.85m (42ft. 2 in.) Maximumtakeoffweight 203,000kg(447,500lbs.) Maximumfuel capacity 95,000kg (209,440lbs.) Estimatedmaximumrange 3000km (1,620nm.) Maximumceiling 19,000m (62,335ft.) MaximumMachnumber 2.4(Envelopeexpandedto 2.0todate) Maximumindicatedairspeed 1000km/hrat 14,500m (540kts at47,500ft.) flight. cations and throttle command information which is used to set the desired thrust through power lever angle Theaerodynamiccontrolsurfacesincludeeight (PLA) in degrees (referred to as throttle alpha by trailingedgeelevons,eachpoweredbytwoactuators, Tupolev). All other engine information including fuel andtwo(upperandlower)ruddersegments. flows and quantities, oil pressures and temperatures, and exhaust gas temperatures are displayed on the flight Propulsion System engineer panel, which is not visible to the pilot. The pilots' throttles mounted on the center console have a The Tu-144LL was derived from a Tu-144D very high friction level, and in normal situations the model originally equipped with Kolesov RB-36-51A flight engineer sets the thrust as commanded by the engines. Since these engines are not in production and pilot in degrees PLA. Autothrottles are normally used consequently could no longer be supported, newer for approaches and landings. Typical PLA settings are power plants were required for the Tu-144LL modifi- 72 ° for maximum dry power, 115 ° for maximum wet cation. Kuznetsov NK-321 engines rated at 55,000 lb power (afterburner), 100 ° for Mach 2.0 cruise, and 59 ° sea level static thrust in afterburner and 31,000 lb dry for supersonic deceleration and initial descent. For take- thrust were selected. These engines are 1.5 meters off weights less than 160 metric tons, 72 ° PLA is com- longer and over 10 mm wider than the RD-36-51A en- manded; for weights from 160 to 180 metric tons, 98 ° gines which necessitated extensive modifications to the PLA is commanded; and for takeoff weights greater engine nacelles and nozzle assemblies. The NK-321 than 180 metric tons, 115 ° is used. Operations in the engines were mounted 1.5 m further forward in the na- 88 ° to 95 ° PLA range are avoided for undisclosed rea- celles, and to accommodate the larger nozzles, the in- sons. board elevons were modified. New higher capacity fuel pumps (jet pumps) were installed in all of the fuel tanks A two-channel autothrottle (A/T) system is avail- with peak pressure capacity of 20 atm. able for approach and landing. It is characterized by a 20 sec time constant and an accuracy of_+7 kin/hr. The The axisymmetric, afterbuming, three stage fan, A/T control panel is located on the center console with five stage intermediate, and seven stage high pressure the two channel selectors, a left/right airspeed command compressor NK-321 engines are digitally controlled, selector switch and a rocker switch to set the speed bug and this dictated a redesigned flight engineer's panel on the respective pilot's airspeed indicator. A ttu'ottle containing eight rows of electronic engine parameter force of 20 kg is needed to override the A/T, or indi- displays. The fuel control consists of a two channel digi- vidual A/Ts can be deselected by microswitches located tal electronic control and a back-up hydromechanical in each throttle knob. If two or more are deselected, the control. The pilot is only presented with N1 RPM indi-

4 entiresystemisdisconnected.Forthesystemtobeen- used to maintain the proper center of gravity (CG) lo- gaged,theflight engineermustengageA/T clutches cation through high capacity fuel transfer pumps. These on theflight engineerthrottlequadrant.TheA/T can transfer pumps are hydraulically driven and controlled beusedfrom160km/hrupto400km/hrindicatedair- by DC power. Fuel boost pumps located in each tank speednormallyorup to 500 km/hr under test condi- are powered by the main AC electrical systems. Tank tions. Use of A/T was authorized only below 1000 m system 4 consists of 6 total tanks, four of which pro- above sea level. vide fuel directly to the engines. A crossfeed capability exists to control lateral balancing. Emergency fuel The variable geometry inlets are rectangular in dumping can be accomplished from all fuel tanks. All shape with a moderate fore-to-aft rake. An internal hori- fuel system information is displayed on the flight engi- zontal ramp varies from an up position at speeds below neer panel, and all fuel system controls are accessible Mach 1.25 to full down at Mach 2.0. Three shocks are only to the flight engineer. produced in the inlet during supersonic flight in order to slow the inlet flow to subsonic speeds. The inlets Numerous fuel quantity probes are used to pro- showed no tendency for stalling or other undesired re- vide individual tank system quantity indications and to sponses during supersonic flight. Full rudder deflec- provide inputs to the CG indicator on the flight engi- tion steady heading sideslip maneuvers were flown at neer panel. A computer within the CG indicator system Mach 2.0 as well as 30 ° banked turns and moderately continuously calculates and displays the CG location. aggressive pitch captures with no abnormal results from The hydraulic fuel transfer pumps, operated manually engines or inlets. Afterburner is required to maintain by the flight engineer, provide fuel balancing using the Mach 2.0 cruise at cruise altitudes. It appeared, but was following transfer routings: to move the CG aft, fuel is not confirmed, that the RD-36-51A engines did not re- pumped from tank 1 to tanks 4, 6, or 8 and fuel is quire afterburner during supersonic cruise even though pumped from tank 2 to tank 8. To move the CG for- the sea level static thrust rating was lower than the ward, fuel can be pumped from tank 8 to tanks 1, 2, 4, NK-321 engines. or 6. The general arrangement of the fuel tanks is shown in figure 3. A discussion of CG management may be Fuel System found later in this report.

The fuel system is comprised of 8 fuel storage Hydraulic System areas composed of 17 separate tanks containing a total operating capacity of 95,000 kg. The nomenclature re- The Tu-144LL utilizes four separate hydraulic fers to fuel tanks 1 through 8, but only tanks 6, 7, and 8 systems, each pressurized by two pumps driven by sepa- are single units. Tanks 1, 2, and 8 are balance tanks rate engines, all of which are connected to separate flight control systems. The flight controls consist of four elevons per wing and an upper and lower rudder. Each control surface has two actuators with two hydraulic {7]U channels per actuator so that each hydraulic system partially powers each control surface. Up to two hy- 7r5 draulic systems can totally fail without adversely af- fecting flight control capability.

{i3 Ci? The four hydraulic systems are powered by variable displacement engine driven pumps. There are no electrically powered pumps. Engine numbers 1 and 2 each power both the number 1 and 2 hydraulic systems [ r 1 and engine numbers 3 and 4 each power both the number 3 and 4 hydraulic systems. Systems 1 and 2 and systems 3 and 4 share reservoirs, but dividers in each reservoir preclude a leak in one system from depleting Figure 3. Fuel tank arrangement

5 the other. Reservoir head pressure is maintained at shaping appears to be parabolic, and this allows pre- 3.2 atm of dry nitrogen. Should nitrogen head pressure cise control at taxi speeds. The 8 ° mode is used for be lost, air conditioning system pressure is utilized to takeoff and landing. The two modes are selected by a provide head pressure. Hydraulic fluid temperature is switch on the overhead instrument panel. maintained within limits by a hydraulic fluid/fuel heat exchanger. The heat exchanger is utilized automatically Electrical System if temperatures exceed 60 ° C. System pressure is nominally between 200 and 220 atm, and a warning indica- The Tu-144 is supplied with main AC power at tion is displayed to the pilot should the pressure in a 115 volts and 400 Hz, secondary AC power at 36 volts system fall below 100 atm. In the event of the loss of and 400 Hz, and DC power at 27 volts. Each engine is two hydraulic systems, an emergency hydraulic sys- connected to its respective Integrated Drive Generator tem is available powered by an Auxiliary Power Unit (IDG). Each of the four AC generators is rated at (APU) air driven pump (or external pneumatic source), 120 kV-A and provides independent AC power to its but the APU can only be operated below 5 km altitude respective bus. There is no parallel generator operation (and cannot be started above 3 km altitude). For emer- under normal circumstances. Each IDG is managed by gency operation of the landing gear (lowering only), a a Generator Control Unit to maintain quality of the nitrogen system serviced to 150 atm is available. If one power supply. Additionally, there are left and right Elec- hydraulic system fails, the aircraft should be slowed to trical Generator Logic Units for power control. Most subsonic speeds. If a second system fails, the aircraft systems can be powered from more than one bus, and should be landed as soon as possible. one generator can provide all of the electrical power requirements except for the canard and inlet anti-ice. The wheel brake system is normally powered by Also available are a separate APU generator rated at the number 1 hydraulic system, but there is a capabil- 60 kV-A at 400 HZ and provisions for external AC ity to interconnect to the number 2 hydraulic system if power. The many fuel tank boost pumps are the main necessary. There is an emergency braking capability electrical power consumers. The high capacity fuel using nitrogen gas pressurized to 100 atm. Indepen- transfer pumps are hydraulically driven and controlled dent braking levers on both the pilot and co-pilot's for- with DC power. Other important electrically driven ward center console areas allow differential braking systems are the canard and the retractable nose. with this system. 36 volt AC power is provided by two main and A locked wheel protection circuit prevents appli- one back-up transformer. This power is used for the cation of the brakes airborne above 180 km/hr airspeed. aircraft's flight instruments, and the total draw is typi- On the ground full brake pressure is available 1.5 sec cally on the order of 1 kW out of a normal 200 kV-A after full pedal pressure is applied. Above 180 km/hr main AC load. on the ground the brake pressure is reduced to 70 atm. Below 180 km/hr brake pressure is increased to 80 atm. The DC system consists of four transformer/rec- tifiers and four batteries. The normal DC load is 12 kW. After the landing gear is retracted, wheel brake pressure at 45 atm is applied to stop wheel rotation. A park- The APU is started from battery power, and DC power ing brake, referred to as a starting brake, is available to is used for communication units, relays, and signaling devices. hold the aircraft in position during engine runups. It is electrically controlled by the pilot and is pressurized to 210 atm. An Essential Bus is supplied by the aircraft's batteries and provides power to an inverter for driving es- A nose gear steering system is available in two sential flight instruments. In an emergency the APU modes of operation. In the high ratio mode, _+60° of may be used to supply 115 volt, 400 Hz AC power. nose gear deflection is available for slow speed taxi- When operating on Essential Power, the normally elec- ing. In the low ratio mode, nose gear deflection is lim- trically driven nose can be lowered with a nitrogen ited to _+8° . Steering is accomplished from either pilot backup system. position through rudder pedal deflection. The pedal

6 Fire Detection and Extinguishing Systems branches. Any one branch can sustain pressurization during high altitude operations. Number 1 and 2 en- Fire detection sensors and extinguishing agents gines and number 3 and 4 engines share common ducts are available for all engines, the APU, and the two cargo for their respective bleed air. The fight system provides compartments. The extinguishing agent is contained in conditioned air to the cockpit and forward cabin areas, six canisters of eight liter capacity each. These canis- and the left system furnishes conditioned air to the ters are divided into three stages. The first stage oper- middle and aft passenger cabin areas. The pressuriza- ates entirely automatically and consists of two of the tion system provides a 15 kg per person per hour air canisters. The remaining two stages are manually con- exchange rate, and the total air capacity is four metric trolled. When an overheat condition is detected, an an- tons per hour. Air is not recirculated back into the cabin. nunciation is displayed on the flight engineer panel The pressurization controller maximum change rate is showing the affected area. The pilot receives only a 0.18 mm Hg per sec. "Fire" light on the forward panel without showing which area is affected. In the case of an APU fire de- Hot engine bleed air is cooled initially to 190 ° C tection, the extinguishing agent is automatically re- by engine inlet bleed air in an air-to-air heat exchanger. leased into the APU compartment. In the case of an The air is then compressed in an air cycle machine engine fire, the pilot can do nothing, since all engine (ACM) to 7.1 atm with an exit temperature of 304 ° C fire extinguishing and shutdown controls are located after which the air is cooled in a secondary heat ex- on the flight engineer panel. changer to 190 ° C or less. If the air temperature is in excess of 90 ° C and fuel temperature is less than 70 ° C Each engine nacelle contains 18 fire detection the air is passed through a fuel-air heat exchanger. Pres- sensors, three to a group. If any one of the groups de- sure at this point is approximately 3 atm. Passage tects an overheat condition, an "Overheat" annuncia- through a water separator precedes entry into the ex- tion is displayed on the flight engineer panel, and a pansion turbine of the ACM. Exit temperature from the first stage canister automatically releases extinguish- turbine must be less than or equal to 30 ° C or the tur- ing agent to the appropriate area. If a second group in bine will shut down. Cockpit and cabin temperature is the nacelle senses an overheat condition, the "Fire" light controlled by the flight engineer using a hot air mix is displayed on the pilot's forward panel. The APU com- valve to control the temperature in the supply ducts. partment has three groups of three sensors each. Any Supply duct temperature must remain between +60 ° group sensing an overheat condition will result in au- and +10 ° C. The nominal engine bleed air pressure is tomatic release of extinguishing agent. The sensors use 5 atm with 7 atm being the maximum allowed before a temperature rate logic for detection of an overheat the engine bleed must be secured. An idle descent from condition. The temperature rate must be 2 ° C per sec high altitude may result in an ACM overheat. In this or greater to indicate an overheat condition. Each sen- case speed must be increased to provide more air for sor has four thennocouples to detect the temperature the inlet air heat exchanger. There are four outflow gradient. With a valid overheat detection, a signal is valves on the left side of the fuselage and two on the sent to the pyrotechnic initiator and valve for the ap- right. The landing gear and brakes are cooled on the propriate canister to discharge automatically. For a sec- ground with air from the outflow valves. ond fire signal the extinguisher must be manually dis- charged by the flight engineer. The flight engineer can The flight engineer controls the air conditioning reset the system to regain the automatic function by and pressurization system. Desired cabin pressure is waiting ten seconds and closing the extinguisher valve set in millimeters of Mercury with 660 mm Hg nomi- to the affected engine. The first stage may be operated nally being set on the ground. During high altitude with battery power only. cruise the ambient cabin altitude is nominally 2800 to 3000 m. Wamings are displayed in the cockpit for cabin Air Conditioning and Pressurization Systems altitudes in excess of 3250 m, and 4000 m is the maximum limit. Air is bled from the cabin in order to cool The air conditioning and pressurization system the flight instruments. There is a maximum tempera- consists of identical, independent left and right ture limit of 30 ° C in the instrument and cargo areas.

7 Anti-Ice System puter. The mutually independent INS units provide attitude and true heading information to the modified There is no provision for wing leading edge anti- Sperry attitude director indicators (ADD and horizon- icing. Flight testing of the Tu-144 prototype indicated tal situation indicators provided to each pilot. Number this was not necessary due to the high speeds normally 3 INS provides inputs to the pilot's instruments, num- flown by the aircraft and the large degree of leading ber 2 to the co-pilot's instruments, and number 1 can edge sweep. The canard, however, is electrically heated be selected by either pilot if necessary. The sensitive for anti-ice protection requiring 20 kV-A of AC power. pitch angle indicator (SPI, figure 4) mounted above the No information was available on engine anti-icing, but center glareshield on the center windshield post is driven the inlets are electrically heated for anti-ice protection. by the number 3 INS. If the navigation computer fails, the pilot can select raw INS data. Each INS can only Navigation and Communications accept 20 waypoints. When within 100 km of the base airport, magnetic heading is used, but outside of that Communication capability consists of standard distance, true heading is manually selected. The crew UHF and VHF band radios and an Interphone Com- has the ability to correct the computed position of each munication System (ICS). Each cockpit crewmember INS separately in 1.6 km increments. can control his communication selection with an ICS control panel. The aircraft is equipped with two VHF Providing guidance to the pilot for the rather com- and one UHF radios, and up to two radios can be se- plex climb and acceleration to cruise conditions and lected for monitoring at one time using a microphone descent and deceleration from cruise conditions is the select wafer switch (which automatically selects the Vertical Regime Indicator (VRI, figure 5). This effec- associated receiver) and a receiver select wafer switch. tive and unique instrument is mounted on each pilot's A variety of aural tones and messages are available in- instrument panel and consists of a horizontal indicated cluding master warning messages, radio altitude calls airspeed display superimposed over a moving vertical (inoperative on the Tu-144LL), and marker beacon profile graphical display. The movable display is driven tones. The annunciation is in a synthetic female voice. by altitude inputs to display the various climb and descent profiles versus altitude. The indicated airspeed Navigation capability consists of three Inertial pointer index travels back and forth on the airspeed Navigation Systems (INS), VOR/DME and ILS receiv- scale, and by adjusting pitch attitude to keep the air- ers, and a Russian version of TACAN. (The ILS re- ceivers are not compatible with Western transmitters.) Desired velocity/altitude Altitude scale The three INS units are controlled by a navigation corn- trajectories (km)\

16-- 15-- [_--__1/_ m /_ VerT2ipa' _____7 14-- 13-- 12-- Airspeed}/' // _k p1 11-- _m,. 10-- 9-- 8-- Pointer .__ 7-- 6-- 500 600 700 800 900 1000 5-- \ 4-- 3-- 2-- 1-- 0-- \ _1-- -2-- ' Airspeed Scale -3-- (km/hr) -4-- Figure 5. Vertical Regime Indicator (VRI) schematic Figure 4. Sensitive Pitch Angle Indicator (SPI) schematic displaying information typical of flight conditions just showing a pitch attitude of +9.5 degrees after takeoff

8 speed index over the appropriate climb/descent curve, Manual Flight Control System the pilot is able to fly the proper profile. The sensitive head-up pitch angle indicator is used in conjunction A schematic of the Tu-144 flight control system with the VRI to maintain the appropriate pitch angle. is shown in figure 6. The system provides a conventional aircraft response with stability augmentation and an aileron-to-rudder interconnect.

2.5 Hz structural filter

pitch rate 1 0.0637s+1 column deg/sec pitch displacement command

2.5 Hz structural filter roll rate column gearing 1 0.0637s + 1 wheel deg/sec displacement rollcommand ID-

wheel gearing _'f 0.15s1 + 1

[_--I_canard or ] aileron-to-rudder

s+l .__.__ extc ,,_ - interconnect

rudder pedal rudder command yaw displacement +i fcommand

pedalgearing

yaw rate ,.. F_Eq._ M _<0.9 + 3.5S deg/sec_ "b_ 3.5s + 1 +

canard or LG ext

sideslip _M > 1.6 f _X_ 0.3s + 1 deg

Figure 6. Tu-144 manual control system

9 8< deg (T.E. down) 20 10 10

7 m m 7

(left) 8a, deg (right) 30 20 10 10 2O 30 I I I I I I 2-- --2

18 8

.---/--18

2 2

Figure 7. Elevon mixer deflection limits

As shown in figure 6, pitch and roll rate sensor lag filters. feedbacks pass through a 2.5 Hz structural filter to re- move aeroservoelastic inputs from the rate signals. Side- The pitch inceptor, or column, force-displacement slip angle feedback is used to improve directional sta- characteristics were depicted on a chart shown to the bility above Mach 1.6 or whenever the canard or land- evaluation team by Tupolev employees. It has been re- ing gear is extended. A yaw rate sensor signal is fed produced in figure 8 as accurately as possible, but some back through a lead-lag filter to allow for steady turn information may have been lost. The deflection of the rates while opposing random yaw motion. column is given in millimeters, and the pull/push force is in kilograms. The feel characteristics are not sym- The pitch and roll command signals are fed metric, with more travel available in the forward, or through mixer logic which limits the combined pitch nose down, direction, as shown in the figure. The exact and roll commands to allowable elevon travel, as shown magnitudes of the aft-most travel forces were not re- in figure 7. Aileron deflection, 8a, represents a differ- corded exactly but are similar to the quantities shown. ential signal subtracted from the symmetric elevator deflection, Be, signal to obtain the right elevon com- Figure 9 depicts the approximate gearing relation- mand; similarly, 8a is added to 8e to obtain the left ships between column displacement in millimeters and elevon command. pitch input to the control system in degrees. Note that the gearing changes depending on whether the landing An aileron-to-rudder interconnect exists to pro- gear or canard is extended. Some data may be missing. vided additional coordination in banking maneuvers between Mach 0.9 and 1.6, and whenever the canard or The roll inceptor, or wheel, force-displacement landing gear is extended, through separate first-order characteristics (as presented by Tupolev) are shown in

10 kg (push) 75. 677 55 -- 235 50.

25. 27--

mm (aft) J 13 200 150 100 50 2.5-- I I I I I I I I 2.5 50 100 150 200 250 mm (forward)

75 kg (pull)

Figure 8. Approximate pitch inceptor (column) force-displacement characteristics

elevon, deg (T.E. down) 30"

Curves are approximate based on sketch of original document canard or gear 22 -- extended

80 14--

10" 10 10 235

8O 175 mm (aft) canard and gear 200 150 100 retracted I I I I I 5O 50 100 150 200 250

mm (forward) 175 10

2O elevon, deg (T.E. up)

Figure 9. Approximate pitch inceptor (column) command gearing

11 35 --/ force, kg (roll right) /I Curves are approximate 30 canard and gear /I based on sketch of original document

retracted _-_&-/ I

extended

10 _0 canard or gear

deg (left) 3-- 1O0 75 50 25 I I I I I I I I 25 50 75 1O0 deg (right)

canard or gear 60 d__ 10 3 extende

I / canard and gear

I/__ retracted force, kg (roll left) 35 Figure 10. Approximate roll inceptor (wheel) force-displacement characteristics

elevon, deg (roll right) Curve is approximate 30 based on sketch of original document

20 18 o

8.5--

10" _ deg (left) IO0 75 50 25 I I I I f I I I I 25 50 75 lOO

deg (right)

8.5-- 10

3O eleven, deg (roll left)

Figure 11. Approximate roll inceptor (wheel) command gearing

12 force, kg (yaw right)

12 100 canard and gear //1-1

retracted _C/ 60

50 o-7°

mm (left) ca _rtd°dgdea r 100 75 50 25 8.5 -- i i i i i i i i 8.5 25 50 75 100 canard or gear _-- mm (right)

extende_...... ,....._J

1 andegear

100 I 6O

force, kg (yaw left)

Figure 12. Approximate yaw inceptor (rudder pedal) force-displacement characteristics

rudder, deg (T.E. right) Curve is approximate 15 15 based on sketch of original document

100

5O mm (left) 1O0 75 50 25 i i i i i 25 50 75 100

mm (right)

100

15 15 rudder, deg (T.E. left)

Figure 13. Approximate yaw inceptor (rudder pedal) command gearing

13 figure 10. This is a symmetric curve that changes the force characteristics if the landing gear or canard is extended. Figure 11 shows the gearing between wheel displacement in degrees and roll stick command input in degrees to the control system.

The yaw inceptor, or rudder pedal, force-displacement characteristics (as presented by Tupolev) are shown in figure 12. This is a symmetric curve that changes the force characteristics if the landing gear or canard is extended. Figure 13 shows the gearing between rudder pedal displacement in millimeters and rudder pedal command input in degrees to the control system.

Autopilot System

The autopilot uses the same servoactuators as the Figure 14. Tu-144 main landing gear manual flight control system and is considered a sub- system of the entire flight control system. The rate Landing Gear dampers in all three axes must be operative for the au- The Tu-144 has a conventional tricycle arrange- topilot to be used. It is a simple two axis system that is ment with twin nose gear wheels and a left and right operated from mode control panels (MCP) located on main landing gear with eight wheels each. The main the pilots' control wheels. Autopilot longitudinal and landing gear, shown in figure 14, has several unique lateral modes include attitude hold, altitude hold, Mach features. hold, bank angle hold, heading hold, localizer tracking, and glideslope tracking. Each mode is selected by Each main gear is a single strut with a dual-twin- pressing a labeled button on the MCR Selection logic tandem wheel configuration. The main landing gear is as follows: For Mach hold to be engaged, attitude includes a ground lock feature that prevents the strut hold must first be selected. Similarly, attitude hold must from pivoting about the bogey when on the ground. first be selected and in operation as indicated by a light This provides the aircraft with a slightly farther-aft on the overhead panel prior to engaging bank angle ground rotation point (with the wheel bogey pivot hold. Two autopilot disconnect switches are located on locked, the aircraft will pitch around the aft wheels in- each MCP, the left one to engage/disconnect the lateral stead of the strut pivot point) as shown in figure 15. channel and the right one to engage/disconnect the lon- This assists in preventing the aircraft from tilting back gitudinal channel. In addition a red emergency discon- onto the tail during loading. nect switch is located on each control wheel. The autopilot channels can also be manually overridden and will disconnect with a 30 mm pitch input or a 15° roll input, respectively. towards nose

Altitude hold can be selected above 400 m altitude, but cannot be used between 0.85 indicated Mach number (IMN) and 1.2 IMN. The lateral modes of the autopilot will command roll angles up to 30 °, but 25 ° is the nominal limit. The longitudinal modes operate between 30 ° nose up to 11° nose down and have a 10 ° elevon trim range capability. - Iockable pivo

Figure 15. Schematic of bogey rotation lock operation

14 lots, as shown in figures 17 and 18. The flight engineer station is located aft of the other three crewmembers on the fight side of the cockpit, as shown in figure 19. The pilots have duplicated round-dial type flight instruments. A set of throttle levers is located between the two pilots, and another set of throttle levers is located at the flight engineer station. Radio and naviga-

.... _ L__-, tion controls are located on an overhead panel that is hinged to swing down in front of the navigator...... 1...... t ------L I I When the hinged nose is raised to cruise position, Figure 16. Schematic of initial gear retraction motion forward visibility is severely limited. To provide atti- Another unique feature of the gear concerns the tude reference information, each pilot has a conven- lateral placement of the main gear attachment and stow- tional hemispheric attitude indicator located in the cen- age when retracted. The Tu- 144 prototype design placed ter of the instrument panel. Because pitch attitude is the four main engines closely together between the land- quite critical for this aircraft, the SPI is located above ing gear struts. In the later models, including the -LL the glareshield directly in the center of the cockpit. This derivative, the engines were moved farther apart. This gauge has an expanded vertical scale with a pointer placed the main landing gear squarely in the middle of indicating aircraft pitch attitude and is calibrated in de- the engine inlet ducting. To accommodate this change, grees. the retracted landing gear was designed to fit between Power lever angles are annunciated on four verti- the inlet ducts of two engines. This was achieved by cal tape indicators in front of the throttle levers. After- having the gear bogey rotate ninety degrees about the burner selection is made above 72 ° PLA, without any strut longitudinal axis (as shown in figure 16) before detent in the motion of the throttle levers to indicate retracting into the tall but narrow wheel well. afterburner selection.

Cockpit Layout All engine controls and displays are contained at The cockpit crew of the Tu-144, including the - the flight engineer's station. The pilots have no direct LL version, consists of two pilots, a navigator, and a knowledge of engine conditions, fuel flow, or power flight engineer. The pilots sit side-by-side, with the setting, aside from the power lever angle and N1 RPM indicators. navigator located on a seat between and behind the pi-

Crew Arrangement

During the three U.S. piloted evaluation flights, the left pilot seat was occupied by the Tupolev chief test pilot, Mr. Borisov. The U.S. evaluation pilot sat in the right pilot seat. The navigator and flight engineer were Tupolev personnel (Mr. Pedos and Mr. Kriulin, respectively).

Flight Equipment

Flight crew were issued partial pressure suits and hehnets with oxygen masks. Parachutes were provided in the aircraft for emergency bail-out. Figure 17. Tu-144LL cockpit seating arrangement

15 Figure 18. Tu-144LL cockpit instrumentation

Figure 19. Tu-144LL Flight engineer's panel

16 Flight Test Planning & Preparation vers consisted of a sharp rap on each control inceptor in an attempt to excite and observe the aeroservoelastic Method of Tests reponse of the aircraft structure.

Through discussion among the U.S. team a con- Due to the lack of simulator support and experi- sensus was reached on what were the highest priorities ence flying the airplane, it was decided not to collect for the evaluation program. It was strongly desired that handling qualities ratings. Such ratings could be mis- both U.S. pilots evaluate the Mach 2 flight regime and leading since they might reflect to a large degree the the approach characteristics to an altitude as low as learning curve with no possibility for repeats to elimi- Tupolev would allow. To assist in the evaluations, spe- nate the effect. Thus, the primary data collected from cifically defined maneuvers were established and re- these flights is pilot comment data. However, during peated for different flight conditions and aircraft con- the post-flight interviews a determination of flying figurations. A brief description of these maneuvers is qualities levels was made. listed below with a fuller description in Appendix A. Planning Process Integrated Test Block (ITB) - This was a standard block of maneuvers designed to provide a consistent Tupolev imposed various requirements which af- evaluation of the airplane for different flight regimes fected the flight test planning. The first flight was to be subsonic and flown with an all Russian crew. The re- and configurations. The maneuvers consisted of: pitch attitude captures, bank captures, heading captures, maining three flights, with the first restricted to sub- steady heading sideslips, and a level deceleration/ac- sonic, were to be flown by one U.S. pilot per flight. celeration. Switching pilots in the middle of the flight was not allowed, as it necessitated the flight engineer and navi- In addition to the ITB, other individual maneu- gator leaving their stations, momentarily leaving flight vers were performed at specific conditions during the critical systems unattended. This meant one U.S. pilot flights: would have two flights, one subsonic and one supersonic, and the other would have one supersonic flight. Parameter Identification (PID) Maneuvers - These However, Tupolev agreed to allow the second U.S. pi- maneuvers involved generating either a sinusoidal fre- lot not flying to observe the second, subsonic flight from quency sweep or a timed pulse train as an input to the the cockpit. Tupolev also preferred to not perform axis of interest. touch-and-goes, which meant that only one approach to touchdown could be done per flight. Simulated Engine Failure - This maneuver involved retarding an outboard engine to its minimum In addition to these requirements there were sev- throttle setting, stabilizing flight, and then performing eral operational restrictions as well. Even though the a heading capture. airplane can take off with a 200 metric tons gross weight, Tupolev recommended against perforating fly- Slow Flight - This maneuver involved pulling ing handling qualities tests at such high weights. A take- back on the column to achieve a certain deceleration to off weight limit of 180 metric tons was therefore ob- capture the minimum speed. This maneuver was done served. In order to provide the correct fuel transfer ca- for level flight and banked flight. pability for the Mach 2.0 flights, fuel had to be consumed until gross weight was below 140 tons prior to Approaches & Landing - Approaches for differ- supersonic deceleration and descent. Approaches with ing configurations were designed such as: canard re- go-around could not be flown until weight was below tracted, lateral offset, manual throttle, nose 0 °, visual, 135 tons, and the maximum landing weight was engine out, and ILS. ILS approaches involved only the 125 tons. A 14 metric ton fuel reserve was planned for localizer, as the airfield glideslope transmitter was in- each flight leading to a target landing weight of operative. 117 tons.

Structural Excitation Maneuvers - These maneu-

17 ; ; ; ; I ; I ; ; ; Mach Number ; Vladimir Sysoev, and the Tupolev pilot, Sergei Borisov, _o---"...... "...... "....___;.... -2,._.....J ,.,,------i.....i--- would review the proposal and offer suggestions relating to safety, efficiency, and feasibility of the profile ...... and each maneuver. Generally two to three days were ---i...... i...... i.... 118- \i--- required to reach consensus on how each flight would ---==--T--==---I---==--1.6-==--_be flown and how much fuel would be loaded on the ---i ...... ,i..... i.4 ---i aircraft. From this consensus Mr. Sysoev would gener- ate a detailed report of the profile and the maneuvers. ___i___'___L__1-2-i- "-_ _i-'_----"\ ! The actual profiles flown (see figures 23 to 25 in Flight t Test Summaries section) were close to planned, with E ---i...... additional approaches added for conservative fuel esti- inates. N_ ,o ...... --i--- < io.7 i : --i--- Flight Readiness Review and Flight Task ...... !ii --i---Examination 5 == iX ==\======On the afternoon before a flight, the various Tu- 144 specialists would meet in a conference room to be briefed on the plan of flight. The aircraft crew would be present as well. Weather and aircraft maintenance ____ -/--_-_ ....i__J___i__J___i___I___i___status would be reported, and the chief of the Calcula- o tion and Experimental Branch, Mr. Sysoev, would 300 400 500 600 700 800 900 1ooo 1100 present an overview of the planned flight maneuvers. Velocity, km/hr The specialists then had an opportunity to raise safety Figure 20. Tu-144 climb and descent profiles concerns and offer modifications to the flight plan. Typically none of the specialists would raise concerns The aircraft center-of-gravity had to be maintained except for the engine specialist, who for flights 22 and at all times within a defined envelope as a function of 23 placed a restriction on a small range of throttle set- Mach number and altitude. The CG was required to be tings for the #2 engine. Once agreement was reached, forward during subsonic operations, and aft for super- the schedule for the next day would be announced, the sonic operations. When accelerating or decelerating flight plan was signed by all parties, and the meeting through transonic conditions, CG had to be maintained concluded in about thirty minutes. within a narrow location only 0.5% mean aerodynamic chord in length. CG location was controlled by fuel Pilot Briefing transfers between forward and aft fuel tanks controlled by the flight engineer. Following the Flight Readiness Review and Flight Task Examination meeting, the aircraft commander, Mr. The climb and descent profiles utilized by Tupolev Borisov, the evaluation pilot, and the U.S. and Russian to climb to and descend from supersonic conditions are engineers would have a meeting to review the maneu- shown in figure 20. These profiles were followed in vers in detail. Final decisions on how the maneuvers flights 22 and 23 up to Mach 2.0 conditions. would be performed were made and any necessary The planning process with Tupolev personnel for modifications based on the flight readiness review were each flight involved joint development of a flight pro- incorporated with the aircraft commander making the file in which the altitude, velocity, fuel loading and CG final decision regarding in-flight issues. position within the above constraints were specified Flight Monitoring and Control along with a sequence of maneuvers. The U.S. team would propose a set of maneuvers and flight condi- The flights were monitored from a control room tions for each flight, usually no more than three days located at the Gromov Russian Federation State Scien- before each flight. The Tupolev flight test engineer, tific Center. Telemetry and radio communications from

18 calreadoutswereprovidedin fivedifferentformats totheengineers,includingatakeoff/landingdisplay, ageneralcontrolsdisplay,apitchdisplay,alateral/ directionaldisplay,andatransonicflightdisplay.In addition,anotherdisplayshowed:thepositionof the aircraftrelativetotheairport,ahorizontalprofileof altitudeversustime,andaltitudeversusairspeedwith theflight envelopeoverplotted.Duringlandingap- proaches, these displays were replaced with a plot of the altitude of the aircraft versus time. In addition to monitoring the progress of the flight on the ground, instrumented fuel quantities were compared to pre- dicted values at points along the profile. Figure21.Gromovflighttestmonitoringarea Following each flight, hardcopy printouts of vari- theaircraftwerereceivedandprocessedat thisloca- ous parameters were plotted using a multicolor pen plot- tion.Fourconsoleswereprovidedtomonitorpropul- ter on B-size graph paper. This allowed many param- sionandfight parameters(figure21).A flightcontrol- eters to be plotted on a single piece of paper and as- ler maintainedcommunicationswith theflight crew, sisted in preparing for later flights. As an example (fig- andfor twoof thethreeU.S.flights,aU.S.flight test ure 22), a chart with 27 parameters overplotted in three engineerwasprovidedwithtwo-waycommunicationscolors clearly showed the change in vehicle weight, withtheU.S.evaluationpilot.Severalparametergraphi- throttle settings, and Mach number that allowed rapid calculation of fuel flow vs. flight conditions.

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19 Flight Test Summaries

Flight 21 - September 15, 1998

Gross weights (metric tons) shown [zero fuel weight 103 metric tons]

10 I Mach 0.9 I 161 155

i o "1o

.{ Q- < Q- = _E "=- _E u_ < "o J o_ o_ e_ m m I I 5m 5E "o <: o_ Q. Q. _< <:

175 166 o > 135 133 124 180 % 0:00 1:00 2:00 3:00 Elapsed Time - hr:min Figure 23. Flight 21 profile flown on September 15, 1998

The flight profile for flight 21, flown on Sept 15, following configurations: a canard retracted configu- 1998, is shown in figure 23. The evaluation pilot was ration using tile ILS localizer, a nominal configuration Gordon Fullerton. Rob Rivers acted as an observer on witll a 100 m offset correction at 140 m altitude, and a tile flight deck. Shortly after take-off a series of ITBs nominal configuration using visual cues. The flight were conducted for the take-off and tile clean configu- ended with a visual approach to touchdown in the nomi- rations at 2 km altitude. An acceleration to 700 km/hr nal configuration. However, due to unusually high was initiated followed by a climb to the subsonic cruise winds the plane landed fight at its crosswind limit, ne- condition of Mach 0.9, altitude 9 kin. Anotller ITB was cessitating tl_e Russian pilot in command to take con- performed followed by evaluations of a simulated en- trol during the landing. Total flight time was approxi- gine failure and slow speed flight. After descent to 2 mately 2 hours 40 minutes. The maximum speed and km evaluations of slow speed flight in tile take-off and altitude was 0.9 Mach and 9 kin. A description of tile landing configurations were conducted, as well as an maneuvers flown is found in Appendix A. A summary ITB and a simulated engine failure in tile landing con- of tile flight written by tile evaluation pilot is found in figuration. Following a descent to pattern altitude tl_ree Appendix B. approaches to 60 m altitude were conducted witll the

20 Flight 22 - September 18, 1998

Gross weights (metric tons) shown 18 157 ITB PID 140 [zero fuel weight 103 metric tons] 16 -2 14 raps, bank captures descent with 12 E

i 10 climb with raps, ll) "0 bank captures -'1 8 <

6 360 ° spiral % z- _ descent o 4 © .__ .d z _ _ > 2 ] 32 129 126 123 120 180 0 0:00 1:00 2:00

Elapsed Time - hr:min Figure 24. Flight 22 profile flown on September 18, 1998

The flight profile for flight 22, flown on Sept. 18, and 9 km altitude. Once again control system raps and 1998, is presented in figure 24. This flight was a super- bank angle captures were performed during the descent. sonic flight with Rob Rivers as the evaluation pilot. At Mach 0.9 an ITB was conducted followed by a de- After take-off a nominal climb profile was flown to scent to pattern altitude. Four approaches were con- establish the supersonic cruise condition of approxi- ducted: a canard retracted approach using the ILS lo- mately Mach 2 at an altitude of 16.5 kin. A series of calizer, a nominal approach (with 0 ° droop nose posi- control system raps and bank angle captures were per- tion on downwind, base, and the initial final legs) with formed during the climb to evaluate lateral/directional an 100 m offset correction at 140 m altitude, and two and structural characteristics throughout the climb pro- visual approaches with the autothrottle off (one with file. At cruise an ITB was performed with the roll and and one without an offset). The flight concluded with a heading capture portions conducted during a 180 ° turn visual approach in the nominal configuration to touch- midway through the cruise portion of the flight. Pa- down. Maximum speed and altitude were Mach 1.97 rameter identification (PID) inputs were conducted and 17 kin. Total flight time was approximately 2 hours, during the remainder of the cruise portion. Upon reach- 10 minutes. A description of the maneuvers flown is ing a minimum fuel weight a nominal descent was con- found in Appendix A. A summary of the flight written ducted to the subsonic cruise condition of 0.9 Mach by the evaluation pilot is found in the Appendix B.

21 Flight 23 - September 24, 1998

I Mach 2.0 I Gross weights (metric tons) shown 18 ITB 138 154 [zero fuel weight 103 metric tons]

nominal descent

E I Mach0.9 I ITB t 10 134 "0 nominal climb

,m 8 <

4 Clean Man Photo Throt 3 eng ILS ILS

2 129 126 123 120 117 179

0 0:00 1:00 2:00

Elapsed Time - hr:min

Figure 25. Flight 23 profile flown on September 24, 1998

The flight profile for flight 23, flown on Sept. 24, tude. A low altitude pass was conducted for photo pur- 1998, is presented in figure 25. This flight was a super- poses with the airplane in a clean configuration fol- sonic flight with Gordon Fullerton as the evaluation lowed by three approaches to 60 m: a visual approach pilot. After take-off a nominal climb profile was flown with the auto-throttle off, a simulated engine out ap- to establish the supersonic cruise condition of approxi- proach and go-around using the ILS localizer, and a mately Mach 2 at an altitude of 16.5 kin. At cruise an nominal approach using the ILS localizer. The flight ITB was evaluated, with the roll and heading capture concluded with an approach to touchdown in the nomi- portions conducted during a 180 ° turn midway through nal configuration using the ILS localizer. Maximum the cruise portion of the flight. One set of frequency speed and altitude were Mach 1.98 and 16.6 kan. Total sweeps was conducted in each axis during the remain- flight time was approximately 2 hours. A description der of the cruise portion. At the end of the cruise por- of the maneuvers flown is found in Appendix A. A sum- tion a simulated engine failure was used to initiate a mary of the flight written by the evaluation pilot is found nominal descent. At 0.9 Mach and 9 km altitude an ITB in Appendix B. was conducted followed by a descent to pattern alti-

22 Observed Vehicle Characteristics gine, were the only engine controls at the pilot station. These power levers move through a range of 0 (idle Ground Handling thrus0 to 115 ° PLA (maximum afterburner). There were no markings on the PLA instrument nor any force de- Nose wheel steering was active at all times on the tent in the throttle quadrant to provide any indication ground and was controlled from either pilot position to the pilot of afterburner ignition. Confirmation of af- by rudder pedal deflection. Two ratios were selectable; terburner operation was a verbal communication from 8 ° and 60 ° of total nose wheel deflection. In the 60 ° the FE. ratio precise control at taxi speeds was easy. A well designed pedal shaping allowed straight ahead control The rerouting of throttle control cables for the without jerking, but pemfitted a very tight 180 ° turn to NK-321 engines resulted in extremely high power le- be accomplished smoothly. The 8° ratio, used for take- ver friction forces. Often the pilot not flying or the flight off and landing, was found to be adequate for lineup engineer manipulated the throttles to ease the workload control throughout takeoff and landing roll, including of the pilot. One technique used was to adjust thrust while landing in a 30-40 kt crosswind. first on two engines and then on the remaining two, because of the difficulty of moving all four at once. Visibility with the nose drooped at 11° (takeoff setting) was adequate for comfortable taxi maneuver- The engines themselves had many operational ing, although it was impossible to see any part of the limits and restrictions, some of which were specific to wing. Accelerations at the cockpit were very mild, con- an individual engine. More time was spent in each pre- sidering its location was considerably ahead of the nose flight readiness meeting reviewing the engine opera- and main gear. The amount of cockpit overshoot re- tion than all other subsystems combined. A 30-minute quired when naming to line up on the runway centefline engine ground run at relatively high power settings was was easily judged. The general feeling during taxi was required to stabilize engine operating temperatures prior much like in the Boeing 747. to flight. If a takeoff was not performed within 1.5 hours of the ground run, the engine ground run had to be re- The hydraulically powered carbon brakes were peated. surprisingly ineffective when cold. The normal pre- takeoff procedure required a brake wammp taxi run. Takeoff/Cleanup Power was advanced to produce a very slow acceleration and the brakes were applied full on. At first there Once lined up on the runway the starting brake was no deceleration at all, but as the brakes warmed was switched on, which applied full hydraulic system they became effective. This procedure was done to be pressure to the brakes. This was sufficient to hold the aircraft while takeoff tt_ust was set and stabilized. Take- ready for a low speed takeoff abort. For landing, warmup was not required because the first brake appli- off thrust setting depended on the gross weight; 115 ° cation at high speeds after touchdown quickly heated PLA (full afterburner) for weights in excess of 180 tons, the surfaces to an effective temperature. 98 ° PLA 0nidrange afterburner) for weights from 160 to 180 tons, and 72 ° PLA 0naximum dry power) for Thrust Management weights less than 160 tons.

Retrofit of the NK-321 engines for the Tu- 144LL Takeoff roll was commenced by releasing the configuration required replacement of the original cock- starting brake and the aircraft accelerated rapidly. For pit engine instrumentation. As in the original design, a 180 ton takeoff weight, V 1 was 255 km/hr, Vr was the flight engineer (FE) station had a complete set of 335 km/hr, Vlof was 355 km/hr, and V2 was 375 kin/ hr. Time from brake release to liftoff was about 30 sec. controls and displays for engine and inlet operation from starnap to shutdown after flight. The pilots had only Directional control presented no problems. Moderate minimal engine information. PLA and N1 were the only back stick pressure produced a slow rotation to a target engine displays, and they were hard to see from the attitude of 8 °. Care had to be taken to not exceed 9° to fight pilot seat. Four power levers, one for each en- preclude contacting the runway with the exhaust

23 nozzles.After liftoff tile pitchattitudehadto be in- pleasant to fly if the aileron feel forces were reduced creasedto approximately16° to controlairspeedbe- by a factor of two. foreraisinggearandflaps.Afterapositiverateofclimb wasestablished,tilelandinggearwasraised. One characteristic became consistently apparent during pitch attitude capture tasks. After reaching a A veryhighambientnoiselevelandamoderate desired attitude and releasing stick pressure to stop tl_e buffetwasexperiencedin tile cockpitwith thenose rate, there was a rebound or drop back in pitch attitude droopedandtilecanarddeployed.Thecanardcouldbe of about one degree. It was necessary to overshoot a seenthroughtile sidewindowin aconstant,obvious pitch target by tiffs amount whetl_er going nose up or vibration.After climbingthrough120m altitude,the down, and whetller at slow speed or supersonic. canardwasretractedandthenosewasraisedtotile0° cruiseposition,resultingin adramaticreductionin the Speed changes resulted in tile expected pitch trim noiseandbuffetlevel. changes of an aircraft witll positive speed stability, and were easily trimmed. The trim rate provided by tile Visibilitywiththenoseretractedwassignificantly wheel mounted electric trim switch was about right. restricted.Theforwardviewwascompletelyblocked, Trim changes due to fuel transfer could be large so it andtheviewtltroughthesomewhatdistortedandcrazed was necessary to continuously trim tl_e aircraft. Canard sidewindowswassopoorthatone'ssensitivitytosmall repositioning produced a very large trim change (de- pitchandevenbankanglechangeswasgreatlyreduced. ploy, nose up; retract, nose down) requiring a constant Controloftl_eaircraftbecameessentiallyaninstrument trim input during tl_e 20 sec extend or retract cycle. task.Quicklyit becameobvioustllattile SPIwasthe bestsourceforpitchattitudeinformation. Response to rudder inputs was conventional, well damped and exhibited a positive dihedral effect. One Handling Qualities could input full rudder, release it completely, and tile aircraft would return to straight flight witll no over- The Tu-144 was equipped witll quad-redundant shoots. As in tile lateral axis, rudder pedal forces were dampers in all tllree axes, all of which were mandated very high. Full pedal deflection required an estimated to be engaged at all times during flight, so the 250 to 300 lb force. unaugmented characteristics of tl_e aircraft were not examined. All of tile above characteristics were invariant witll speed and configuration, from Mach 2 cruise to Pitch control forces were moderate and not un- minimum speed in the landing pattern, witll some small usually heavy for a large aircraft. Small movements of degradation near Mach 1. tl_e control column resulted in significant pitch motion. Initially tllere was a tendency to overcontrol during pitch Acceleration/Climb maneuvers, especially at high subsonic and supersonic speeds. However, pilot compensation for tiffs charac- Once in a clean configuration, tile normal proce- teristic was easy. In contrast, lateral control forces were dure called for setting the throttles to maximum dry very high and wheel deflections required for even low power, 72 ° PLA, and accelerating in a climb to 600 kin/ roll rates were large. One could fly pitch witll one hand hr. At 2 km altitude a furtller acceleration to 700 km/hr but one tended to use two hands when a bank change was accomplished in order to intercept tl_e climb pro- was needed. Control harmony was moderately objec- file depicted on tile Vertical Regime Indicator (VRI). tionable. Due to tiffs deficiency, it was easy to make This mechanical instrument portrayed altitude versus inadvertent pitch inputs on purely lateral tasks. airspeed on a moving tape (driven by actual altitude) and tl_e desired altitude/speed profile to maintain for Roll response was very well damped and tllere optimum climb, normal descent, and emergency de- was no adverse or proverse yaw even witll large lateral scent perfommnce. It was found to be an intuitive pre- inputs and no rudder inputs. Bank angle captures were sentation. At Mach 0.95, when tile CG had been ad- easy to accomplish precisely. Aileron feel forces were justed to 47.5%, tl_e tl_rottles were advanced to full af- extremely heavy; tile aircraft would be much more terburner power (115 ° PLA), and tiffs tllrottle setting

24 wasmaintaineduntilleveloff. To simulate an engine failure, an outboard engine was reduced to tl_e minimum allowable thrust setting Becauseof tl_erestrictedoutsideview,smallbut while at Mach 2. A mild yaw resulted which was easily significantpitchattitudechangeswerenotobviousto controlled witll rudder trim. There were no apparent tilepilotexceptasobservedontheSPI.TheADI was effects on any otl_er engine. The remaining three en- notsensitiveenoughfortightcontrolofpitchattitude. gines were advanced to full afterbuming tllmst but tile Maintainingtiledesiredclimbprofilewasafull time, thrust was insufficient to maintain speed and a slow highworkloadtaskduetotileinherentflightpatllsen- deceleration resulted. sitivity to smallpitch attitudechanges,pooroutside visibility,andtileneedfor frequentpitchtrimchanges Descent duetocenterof lift shiftingandCGadjustmentsastile Machincreased.Also,thelocationoftileinstantaneous Descent from Mach 2.0 cruise was initiated by centerofrotationneartl_ecockpitdeprivedtl_epilotof retarding tile tllrottles from tile afterburner range to 59 ° motioncuesduetopitchrate.Frequentreferencetothe PLA and decelerating to 800 km/hr while descending. SPIwasfoundto beessentialfor smoothcontrolof The VRI descent profile was intercepted during tile climbanddescentprofiles. descent. The next power reduction to 52 ° PLA occurred at 14 km altitude witll a tlfird power reduction to 35 ° Thepitchcontroltaskremaineddifficultuntillev- PLA at 11 kin. At Mach 0.8 the tllrottles could be re- elingoff at Mach2.0 at aninitial cruisealtitudeof duced to idle. From Mach 2.0 and 17 km to Mach 0.9 16.5kin. Fromtakeoffto leveloff at Mach2.0and and 9 km altitude took 7 minutes. The CG had to be 16.5kmtook19minutes. moved aft from 46% to 47.5% MAC prior to passing transonic speeds followed by a rapid forward shift to In orderto examinehandlingqualitiesthrough- 41% MAC for subsonic speeds. The shifting location outtheenvelopetl_eautopilotwasnotusedduringthe of tile CG might have contributed to some of the higher climb,cruiseanddescentphasesof anyflight.It was workload in controlling tl_e pitch attitude in tl_e tran- notauthorizedforusetransonically,betweenMach.85 sonic region. The pitch sensitivity increased in tile Mach andMach1.2.Theautopilotis describedin tile Air- 1.2 to Mach 0.95 region especially near Mach 1.0 witll craftDescriptionsectionof thisreport. a quite definite transonic pitch up just below Mach 1.0. At subsonic speeds tile overall descent task became Throughouttileclimbprofile,smallstepchanges easier. inlateraltrimoccurredfrequently,requiringlateraltrim switchinputs.Noexplanationforthiswasdetermined. The SPI was provided primarily for pitch attitude information in tl_e descent. It was a very useful instru- Supersonic Cruise ment, but its sensitivity, accompanied by a perceived delay between attitude change and flight patll angle After stabilizing at cruise Mach, tile trim changes change, produced a relatively high workload task. This due to speed changes and fuel shifts ceased, reducing was exacerbated by a lack of pitch reference cues witll tile pilot workload considerably. At Mach 2.0 it was the nose retracted and a slight difficulty in getting ad- noticed tlmt a step increase in pitch force gradient oc- equately sensitive pitch reference information from tl_e curred after tile aft column breakout. No nuisance pitch/ ADI. Very small pitch inputs (less tlmn 0.25 cm) re- roll aerodynamic coupling existed, and tile aircraft was sulted in up to two or more degrees of pitch attitude very stable directionally. The engine inlets appeared change. Therefore, tile pilot's control gain had to be insensitive to sideslips (up to 4.5 ° of rudder input) or decreased during tile flight in order to not over-control to pitching motions. the pitch axis. About half of tile starting fuel load (80 tons) was Structural Characteristics consumed for takeoff, climb, and acceleration to Mach 2. The aircraft burned approximately 1 ton of fuel Passing 4 km altitude in tile climb, tile first of a per minute and required midrange afterburner to main- number of control raps for the purpose of exciting tain Mach 2.0.

25 aeroelasticmodeswereaccomplished.Thesemaneu- final on a shallow (2 °) glideslope with established check verswererepeatedat 6 km altitudeandacceleratingpoints of range and altitude. The aircraft was stabilized throughMach0.7,0.9,1.1,1.4,1.8,andatleveloff at in level flight at the final approach speed (on the order 16.5kmatMach2.0.Thecontrolrapsconsistedofrapid of 360 km/hr depending on fuel weight) which resulted stepinputsof smallmagnitudein eachcontrolaxis. in an AOA of 10°. At the proper range from the runway Responsesin allthreeaxesbehavedsimilarlyatallsub- the nose was lowered to 8°, setting up a sink rate that sonicspeeds.Arapid1-2cmforwardoraftpulseof the was held constant all the way to the runway threshold. controlcolumncausedacockpitverticaloscillationof 1to 2 Hzwith 2 to 3overshootsof about1° in pitch Flight at AOA of 10° was definitely on the back attitude.A rapid20 degreewheelinputresultedin a side of the thrust required curve as indicated by fre- lowerfrequencybuthighermagnitudelateralresponse quent power adjustments. To reduce workload the of 1° to2° and4-6overshoots.Rudderrapsresultedin autothrottles were normally used throughout the ap- ahnostpurelydirectionalaeroelasticresponseof 4-6 proach until just prior to touchdown. The autothrottle overshootsataboutthesamefrequencyastheroll re- was a relatively low frequency system (20 sec time sponse. constant) but was effective at holding the desired speed. Manual throttle approaches were flown and were found Theaeroelasticresponseof theaircraftdiffered not to be excessively difficult, except for the physical asthespeedincreasedandalsodependedonwhichaxis workload from having to overcome the extremely high wasexcited.Eachaxisof excitationresultedin adif- friction forces. Flight path angle was controlled with ferentaeroelasticresponse.In particular,theyawand pitch inputs and speed with the throttles. Despite being roll aeroelasticreponsesappearedtobedecoupled.In a back-sided aircraft, the engine throttle response time general,theaircraftseemedmoreheavilydampedsu- constants were reasonable, and airspeed deviations on personicthansubsonic.At Mach2.0in all axesthere- approach were kept within 2 to 3 kin/hr. sponsewasofhigherfrequencyandsmallermagnitude thanatsubsonicspeeds. There was a noticeable pitching moment change with thrust change. If large power corrections were Low Speed Characteristics made, large pitch trim changes followed. Similarly, if large pitch inputs were made, fairly rapid speed changes The total airframe drag increased dramatically as resulted which required throttle adjustment. It was pos- speed was decreased. Slowing to the angle of attack sible to get into a throttle/pitch input coupling situa- (AOA) limit of 16 ° at 9 km altitude, maximum dry tion. Once one learned to direct constant attention to thrust was insufficient to accelerate in level flight. Climb maintaining the proper pitch attitude using the SPI, rate at maximum dry thrust was substantially lower (by control of the aircraft became easier. a factor of four) at 430 km/hr compared to climb rate at 700 kin/hr. Control response remained excellent at mini- Go-arounds were initiated by setting 72 ° PLA mum speeds with motions in all axes well damped. which provided a brisk acceleration. With a positive climb rate the landing gear was raised. Passing 120 m Traffic Pattern altitude, the canard was retracted if desired, while applying the required nose-up trim inputs. Power was Once the nose was lowered to the 17° droop posi- normally reduced passing 300 m to reduce the climb tion for landing, visibility was adequate for approach angle. With the landing gear extended, the bank angle maneuvering. Canard extension had to be countered limit was 15° and an aural voice warning triggered when with several seconds of nose-down trim. Nose-down, that limit was exceeded. With the gear retracted, it was canard-extended flight produced the same noisy buf- comfortable to bank up to 30° . feting cockpit environment encountered after takeoff. Landing gear extension produced no noticeable noise, The first evaluation flight was flown under strong vibration, or trim change. gusty wind conditions with moderate turbulence at pattern altitude. Due to the high lateral forces and less than The Tupolev preferred approach technique was desirable control harmony, the constant control inputs to set up on a long (10 to 12 kan) stabilized straight-in

26 requiredtomaintainattitudebecamephysicallytiring. oscillation(PIO). Whentileautopilotwasengaged,it performedaccept- ablyin timrelativelyturbulentconditions. Landing/Ground Roll

Anengine-outconditionwassimulatedbyretard- Descending tltrough 15 m, ground effect caused ing an outboardengineto idle. At normallanding a strong nose down pitching moment which required a weightstherewasadequatethrustwitl_theremaining firm pull on tl_e control column to maintain attitude. enginesatdrythrustorlessto fly astandardpattern, Normal technique was to maintain or slightly increase approach,andgo-around.It wasfoundduringathree- the pitch attitude, and allow tl_e aircraft to fly onto tl_e enginepatternflownon thefirst missionat aheavier runway. The ground effect cushion provided a soft land- weighttlmtmaximumdrytl_rustwasbarelyadequateing in each case. Care had to be taken to not overflare for amissedapproach.In botl_casestl_erewasplenty or to hold tl_e aircraft off, allowing tl_e pitch attitude to ofruddercontrolpowertomaintaindirectionalcontrol exceed 10° and risk contacting tile engine exhaust atall times.Rudderforcescouldbetrimmedoutwitll nozzles witll tile runway. timelectricruddertrimswitch. Derotation was easily controlled. The long nose Severalapproachesweremadewith tile canard gear and nominal 3.5 ° pitch attitude in tl_e tl_ree-point retractedwhichincreasedtimrequiredfinal approach stance resulted in tl_e appearance of a significant nose- speedbyabout30kin/hr.Therewasaslightreduction up attitude at nose gear touchdown. The drag parachutes in tilebuffetlevelsensedattilecockpitbutnochange were deployed after nose gear touchdown, and wheel in handlingqualitieswasnoticed. brakes were applied below 220 kin/hr. Only light braking was required to stop tim aircraft. A noseretractedapproachwasflownin orderto evaluatetheabilityto landin tiffsveryrestrictedvis- On tim first evaluation flight, strong gusty cross- ibility condition.Forwardvisibility wasahnostnon- winds were encountered at landing. The aircraft was existentduetotimmetalskinontopof thenoseblock- landed witll tile recommended crosswind technique, ingtilepilot'sforwardfield of view.A nose-retractedwings level in a crab. It tended to align itself witll tim landingmaybepossibletoaccomplishthroughtimuse runway after touchdown witl_out much rudder input, of thesidewindowandananglingapproach,buttiffs and tlmre was ahnost no rolloff tendency. The rudder wasnotevaluatedtotouchdown. pedal nose wheel steering in tile 8° ratio was exactly right for maintaining runway centerline during drag Oneclean(gear,nose,andcanardretracted)con- chute deployment and deceleration. figurationlow approachpasswasflown for photo- graphicdocumentation.Afterliningup with tl_erun- Conclusions wayabout10km out,tl_enosewasraised.It wasim- possibleto seeanypartof tl_eaerodromeor its sur- The opportunity to fly tim Russian supersonic roundingstructures,andtl_eonlylineupinformation transport allowed U.S. pilots to observe tim flight char- availablewastheILScoursedeviationindicator. acteristics of one of only two such examples of supersonic passenger-carrying aircraft. From this opportu- Severallateraloffsetapproacheswereflownto nity, several observations are made, and some conclu- examinehandlingqualitiesin tiffshighworkloadtask. sions can be drawn witl_ application to future super- Theoffset was 100m to the right of therunway sonic aircraft of tiffs type. centefline.A lineupcorrectionwasstarteddescending tllrough140m altitude,andamissedapproachwas The Tu-144 exhibited heavy lateral control forces initiatedat 60m. Sincetimapproachescouldnotbe throughout tim flight envelope. The pitch axis con- flowntotouchdown,theworkloadfortiffstaskwasnot trol forces were more appropriate, but still a bit ashighashadbeendesired.Theaircraftresponded heavy, for tiffs class of aircraft. nicelyduringmoderatelyaggressivelateralmaneuver- ing.It waseasytojudgetimcorrection,androll out • The lateral axis had a low roll response sensitivity wasaccomplishedwith notendencyfor pilot-induced at all speeds which required large control wheel de-

27 flectionsto achieveevenmildroll rates.Thepitch 4. Apitchaugmentationsystemthatprovidesautomatic axisdemonstratedacorrespondinglyhighersensi- trimming,suchasaflight-path-angleresponse-type tivity throughouttheflight envelope.Theresulting system,mightreducepilot workloadon futuresu- lackof harmonyin thewheelandcolumncontrol personicaircraft. sensitivityledto instancesof couplingroll inputs intopitchinputs. Otherobservationswere: Duetopowerchanges,fueltransfers,andotheruni- • Theeffectsofaeroelasticitywereapparent,butwere dentifiedlateraltrim changesduringclimb and consideredminoranddidnotaffecttheflyingquali- cruise,continualadjustmentsof bothpitchandroll tiesof theTu-144. controllerscontributedtohighworkloadfor thepilot. Poorvisibility with thenoseraisedandprob- • Makingthepatternturnto finalwaspossiblewith lemswithpitchresponsedynamicsaddedtothedif- thenoseretracted,butlandingwasnotattempted. ficultyof pitchcontrol. Thefield-of-viewwith thenosein theretractedpo- sitionis extremelylimited,buta nose-uplanding Duetothedeltaplanformof thewing,theinstanta- appearstobefeasible. neouscenterof rotationof theTu-144in pitchap- pearedtobenearthecockpit.Thisresultedin little • Thegenerallateral/directionalandlongitudinalchar- normalaccelerationchangewithpitchchangesand, acteristicsremainedfairly constantthroughoutthe consequently,nomotioncuesforthepilot.Thischar- flightenvelope. acteristicis notfoundin conventionalaircraftwith afttails. Goodcrosswindlandingcapabilitywasdemon- stratedonthefirstflight,witha10m/secdirectcross- wind. Excessivethrottlefriction wasnoted,leadingto higherworkloadsduringmanualthrottlelandings. Operationalconcemsincludetheneedtowarmup Normally,Tu-144landingswereperformedwiththe thewheelbrakesduringtaxifortakeoff,therequire- autothrottleengaged.Backside-of-the-power-curvementtowarmuptheenginespriortoflight(thismay characteristicswerenotedbutwere not objec- tionable. beuniqueto theTu-144LL)andtheutilizationof dragchutesonlandingrolloutfordeceleration.Each of theseconcernswouldneedtobemitigatedfor a Eachof thecharacteristicsjustdescribedledto commercialaircraft. anoverallhandlingqualitiesratingof Level1for the roll axisandLevel2 for thepitchaxis(asdefinedby Marginalclimbrateswerenotedwhenusingthree theCooper-HarperPilotRatingscalein reference1). enginesatmaximumdrypowerwhile thevehicle TheLevel2ratingis attributedto poorpitchattitude washeavy;otherwise,controlof thevehicleduring informationandfrequenttrim changesduringclimb/ descent. asymmetricthrustconditionswasnotdifficult. Basedononlylimitedevaluationtime,acanard-re- Lessonsto belearnedfor futuresupersonicair- tractedapproachshowedanobvioustrim andair- craftof thistypeinclude: speeddifferencefromanapproachwith thecanard 1.Controlforcesshouldbereducedin allaxes,includ- extended,butotherdifferencesin handlingqualities ingthrottle. werenotnoticeable. 2. Pitchcontrolsensitivityshouldbeafunctionofflight Inadequateengineinstrumentationwasprovidedat regimetoincreasecontrolharmony. thepilot stations.Changesin powersettingswere notapparentto thepilotsexceptin pitchcoupling 3. An autothrottlesystemmaynotbemandatoryfor dueto alack of visualor auralcues.Particularly manuallandings. missedwasanyindicationof afterburnerignition.

28 • Groundeffectsduringlandingassistedin cushion- ingthevehicleduringthelandingflare.

References

1. Cooper, G. E.; and Harper, R. R, Jr: The Use of Pilot Rating in the Evaluation of Aircraft Handling Qualities. NASA TN D-5153, April 1969.

29 Appendix A: Description of Maneuvers

This appendix describes a typical definition of the maneuvers fown in the evaluation program. Modifica- tions to target values were required for some of the maneuvers based on flight condition. This section is intended to provide a general idea of how the maneuvers were defined.

DecelerationAcceleration

a. Reduce throttle from trim position by approximately 20 °. b. Decelerate and capture an airspeed 70 km/hr less than trim. c. Advance throttle to full dry power setting. d. Accelerate and re-capture original airspeed.

Pitch Attitude Capture

a. From trim attitude pull column back to capture a +3 ° pitch attitude increment. b. Push column forward to capture original trim attitude. c. From trim attitude push column forward to capture a -2 ° pitch attitude increment. d. Pull column back to capture original trim attitude. e. Throughout a.-d. keep normal acceleration between 0.8 and 1.2 g.

Bank Angle Capture

a. From steady level flight apply right wheel to capture a +30 ° bank angle. b. Apply left wheel to capture level flight. c. From steady level flight apply left wheel to capture a -30 ° bank angle. d. Apply right wheel to capture level flight.

Heading Captures

a. From steady level flight apply right wheel to capture a +30 ° bank angle. b. Maintain bank angle and capture +20 ° heading increment. c. Apply left wheel to capture a -30 ° bank angle. d. Maintain bank angle and capture original heading. e. Repeat in opposite direction.

Steady Heading Sideslips

a. From steady level flight apply a series of rudder deflections of +2, +4, +6, and +7.5 °. b. Apply appropriate wheel deflection to maintain constant heading, stabilizing for 5 sec on each rudder deflection. c. Repeat in opposite direction.

Simulated Engine Failure

a. From steady level flight retard throttle #1 to idle. b. Wait 5 sec and then stabilize transient, maintaining a bank angle less than _+5°. c. Advance three remaining throttles to capture original airspeed. d. Perform heading capture maneuver. e. Recover by slowly advancing #1 throttle and re-establish original flight condition.

3O Slow Flight

a. From steady level flight pull column back and establish a 2 km/hr per sec deceleration. b. At minimum airspeed or warning stop deceleration and hold condition for 3 sec. c. Recover by pushing column forward and establishing original flight condition. d. From steady level flight establish a 30 ° bank turn e. Pull column back and establish a 2 km/hr per sec deceleration. f. At minimum airspeed + 10 km/hr or warning, stop deceleration and hold condition for 3 sec. g. Recover by pushing column forward, rolling wings level and establishing original flight condition.

Frequency Sweep PID maneuver

a. Allow 15 sec of steady level flight b. Commence longitudinal sinusoidal input of 1.5 cm to 2.0 cm (no greater than 0.8 to 1.2 g) with a period of oscillation of 20 sec. c. Increase frequency at constant amplitude over 80 sec to a period of 1 sec. d. Wait 15 sec and then recover to original conditions. e. Repeat a.-d. for lateral wheel input of 15 ° to 20 ° (bank angles between 5 ° and 10°). f. Repeat a.-d. for rudder pedal input of 1.5 to 2.0 cm (heading changes between _+5°).

Timed Pulse Train PID maneuver

a. Allow 5 sec of steady level flight. b. Input 2.0 cm forward (of trim) control pulse for 3 sec. c. Input 2.0 cm aft (of trim) control pulse for 2 sec. d. Input 2.0 cm forward (of trim) control pulse for 2 sec. e. Allow 5 sec of steady level flight. f. Input 2.5 cm left pedal for 3 sec. g. Input 2.5 cm right pedal for 2 sec. h. Input 2.5 cm left pedal for 2 sec. i. Return pedal to neutral while inputting 20 ° right wheel. j. Keep right wheel input for 1 sec. k. Input 20 ° left wheel input for 1 sec. 1. Release controls for 10 sec. m. Repeat a.-e. with 4.0 cm amplitude. n. Repeat e.-1. with 40 ° (wheel) and 5.0 cm (pedal) amplitude.

Structural Excitation maneuver

a. Sharply deflect control inceptor for a single axis approximately 1-2 cm (column or pedal) or 10° (wheel). b. Release controller and allow aircraft to aeroelastically respond until all motions are damped

31 Appendix B. Extended narratives for each U.S. piloted flight

Flight 21

Date of Flight: September 15, 1998 Flight Crew: Pilot in Command: Sergei Borisov Evaluation Pilot: Gordon Fullerton

Navigator: Victor Pedos Flight Engineer: Anatoli Kriulin Takeoff Time: 10:58 Local

Landing Time: 13:43 Local Flight Duration: 02:45 Takeoff Weight: 180 metric tons Landing Weight: 124 metric tons Landing Fuel: 21 metric tons Total Fuel Bum: 56 metric tons Takeoff CG: 40.5%

Landing CG: 40.7%

Flight Summary

The engines were started by the flight engineer in a 2, 3, 4, 1 order. After completing a short pre-taxi checklist the parking brakes were released. The aircraft began to roll slowly with the power levers still at idle. With the 60 ° steering ratio selected the runway was entered and a brake warm-up procedure was done, consist- ing of setting the power levers to 35 ° PLA and applying the brakes. At first full brake pedal deflection would not slow the aircraft, but as the carbon brakes wanned they became more effective.

Holding in lineup position, the steering ratio was set to 8 °, and the engines were set at 98 ° PLA (partial afterbumer).The aircraft accelerated rapidly through rotation and liftoff speeds.

After takeoff the landing gear was retracted but the canard and nose were left in takeoff configuration (deployed and 11°, respectively). Leveling at 2000 m a series of maneuvers, called the Integrated Test Block (ITB), was performed to evaluate handling qualities at a heavy weight takeoff configuration. The ITB tasks included a deceleration and acceleration, pitch attitude captures, bank angle captures, heading captures, and steady heading sideslips.

The nose was raised, canard retracted and a climb to subsonic cruise conditions (Mach 0.9 and 9000 m) was made where the ITB series was repeated. Then engine #1 was retarded to near-idle thrust to evaluate handling in an asymmetric thrust condition.

32 Afterreturningtosymmetricthrusttimaircraftwasslowlydeceleratedtoapproachtimangleof attacklimit of 16°, tlmnrecoveredbydecreasingpitchattitudeandincreasingpower.Theslowflight characteristicswere furtl_erinvestigatedbyrepeatingtiffsprocedurein a30° bank,recoveringbylevelingtl_ewingsastl_rustwas increased.

Powerwasreduced,descendingto2000m andconfiguringtheaircraftwitlathenoseat 11° andcanard deployed.Theslowflightproceduresjustdescribedwererepeated. Afterestablishingalandingconfiguration,nose17°, canarddeployed,andlandinggeardown,theITB maneuverswereperformed,followedbyanotllerslowflight investigationwitll wingslevelandin a 15°bank. Next,anenginefailurewassimulatedbyretardingtim#1enginetonearidle.After a5 secdelaywitl_no controlinputsto observetile aircraftresponse,powerfor levelflight wasseton engines2,3,and4, andtile aircraftwastrimmed.Adescentof4m/secwassetupsimulatinganormalapproachglideslope.At 1500m AGL powerwasadvancedtomaximumdrytllruston2,3, and4,establishingashallowclimb.Thelandinggearand canardweretl_enretracted.

Sincegrossweightwasslightlyabovetimmaximumlimit forlanding,thefirst approachwasapassdown tilerunway30centerlineatabout400m altitude.Windswereunusuallystrong,gustingfromtile soutllupto 20m/sec(40knots)witll veryturbulentconditionsatpatternaltitudeandbelow. A closedpatternwasflownleadingtoavisualapproachtorunway30witl_thecanardretractedandinitiat- ingago-aroundat60mabovethesurface. Thenextapproach,alsotorunway30,wasintentionallyalignedabout100mrightoftherunwaycenterline untildescendingtllrough140mwhenanS-turnmaneuverwasdonetocorrecttilelineuperror.Asbefore,ago- aroundwasstartedat60m.

A plannedlowpassdowntilerunwayforgroundeffectsdatawascanceledbecausethewindswerefarin excessof the2.5m/seclimit. Becauseof tl_estrongtailwindcomponentonrunway30,theaircraftwasmaneu- veredtoarighthanddownwindlegforrunway12,followedbyanapproachandgo-aroundat60m. Theairporttrafficareawasdepartedtotimeastouttoadistanceof about60kmtobumdownfuelpriorto timfinallanding.Theautopilotwasengagedandevaluatedduringtiffsdelay. Returningto timrunway12pattern,avisualapproachandfull stoplandingwasmade,in astrongcross- windfromtile right,tile strongesteverencounteredin thisaircraftby Mr. Borisov.Theaircraftwasstopped usingthedragchutesandlightbraking.Afterjettisonof tilechute,tileaircraftwastaxiedbacktotilestartuparea andshutdown.

33 Flight 22

Date of Flight: September 22, 1998 Flight Crew: Pilot in Command: Sergei Borisov Evaluation Pilot: Rob Rivers

Navigator: Victor Pedos Flight Engineer: Anatoli Kriulin Takeoff Time: 11:08 Local

Landing Time: 13:23 Local Flight Duration: 02:15 Takeoff Weight: 184 metric tons Landing Weight: 110 metric tons Landing Fuel: 16 metric tons Total Fuel Bum: 65 metric tons Takeoff CG: 41%

Landing CG: 40.4% Weather:

Scattered clouds at 6 kin, winds 150 degreees at 2-3 m/s, altimeter setting 755 mm Hg QFE

Flight Profile

The flight profile included takeoff and acceleration to 700 kilometers per hour (km/hr) to intercept the climb schedule to 16.5 kilometers (kin) and Mach 2.0. The flight direction was southeast toward the city of Samara on the Volga River at a distance of 700 km from Zhukovsky. Approximately 20 minutes were spent at Mach 2.0 cruise which included an approximately 190 degree course reversal and a cruise climb up to a maximum altitude of 17.3 kin. A descent and deceleration to 9 km and Mach 0.9 was followed by a brief cruise period at that altitude and airspeed prior to descent to the traffic pattern at Zhukovsky Airfield for multiple approaches followed by a full stop landing on Runway 30.

Flight Summary

After all preflight checklists had been completed, the evaluation pilot taxied Tu-144LL Serial Number 77144 onto Runway 12, and the brake bum-in process was accomplished. At 11:08 brakes were released for takeoff, power was set at 98 ° PLA (partial afterburner), the start brake was released, and after a 30 sec takeoff roll, the aircraft lifted off at approximately 355 kin/hr. The landing gear was raised with a positive rate of climb, the canard was retracted out of 120 m altitude, and the nose was raised out of 1000 m altitude. The speed was initially allowed to increase to 600 km/hr and then to 700 km/hr as the Vertical Regime Indicator (VRI) profile

34 wasintercepted.Powerremainedat72° PLA(maximumdrypower)for theclimbuntilMach0.95andCGof 47.5%atwhichpointthethrottleswereadvancedtomaximumpower,115° PLA.Theclimbtaskwasahigh workloadtaskduetothesensitivityof theheaduppitchreferenceindicator,thesensitivityof thepitchaxis,and thecontinualchangein CGrequiringalmostcontinuouslongitudinaltriminputs.Also,sincetheinstantaneous centerofrotationis locatedatthepilot station,therearenocockpitmotioncuesavailabletothepilotfor pitch rateor attitudechanges.Significantpitchratescanbeobservedon thepitchattitudereferenceindicator(SPI) thatarenot sensedby thepilot. Duringtheclimbpassing4 kin, thefirst of arepeatingseriesof bankangle captures(_+15°) andcontrolrapsin all threeaxes(toexciteanyaircraftstructuralmodes)wascompleted.These maneuverswererepeatedat6kmandwhenacceleratingthroughMach0.7,0.9,1.1,1.4,and1.8.Thebankangle capturesdemonstratedratherhighroll forcesandrelativelylargedisplacementsrequiredfor smallroll angles.A welldamped(ahnostdeadbeat)rollmodeatallairspeedsuptoMach2.0wasnoted.Thecontrolrapsshowedin generalahighermagnitudelowerfrequencyresponsein allthreeaxesatsubsonicspeedsandlowermagnitude, higherfrequencyresponsesatsupersonicspeeds.Thepitchresponsewasin generalof loweramplitudeand frequencywithfewerovershoots(2-3)thanthelateralanddirectionalresponses(4-5overshoots)atall speeds. Alsoof interestwasthattheaxisexhibitingtheflexibleresponsewastheaxisthatwasperturbed,i.e.,pitchraps resultedin essentiallyonlypitchresponses.Themotionsdefinitelyseemedtobeaeroservoelasticin nature,and with thestrongdampingin thelateralanddirectionalaxes,normalcontrolinputsresultedin well damped responses. Leveloff at16.5km andMach1.95occurred19minutesaftertakeoff.Theaircraftwasallowedtoaccel- eratetoMach2.0IMN asthethrottleswerereducedto98° PLA,andaseriesof controlrapswasaccomplished. Followingthis,aportionof theIntegratedTestBlock setof maneuversconsistingof pitchcaptures,steady headingsideslips,andaleveldecelerationwascompleted.Thepitchcapturesresultedin slightovershootsand indicatedamoderatedelaybetweenpitchattitudechangesandflight pathanglechanges.Thesteadyheading sideslipsshowedaslightpositivedihedraleffect,butnomorethanapproximately5° angleofbankwasrequired to maintainaconstantheading.No unpleasantcharacteristicswerenoted.At thispointthefirst setof three longitudinalandlateral/directionalparameteridentification(PID)maneuverswerecompletedwith nounusual results.By thistimeacoursereversalwasnecessary,andthebankangleandheadingcaptureportionsof theITB werecompletedduringtheover180°mmwhichtookapproximately7raintocompleteatMach1.95.Duringthe inboundsupersonicleg,twomoresetsof PIDmaneuverswith higheramplitude(doublethefirst set)control inputswerecompletedaswereseveralmoresetsof controlraps.Maximumaltitudeachievedduringthesuper- sonicmaneuveringwas17.3kin. ThedescentanddecelerationfromMach2.0and17kmbeganwithapowerreductionfromthenominal 98°PLAto59° andadecelerationto 800kan/hr.Duringthedescentbankanglecaptures(_+30°) andcontrolraps wereaccomplishedatoraboutMach1.8,1.4,1.1,and0.9with similarresultsasreportedabove.Theaircraft demonstratedincreasedpitchsensitivityin thetransonicregiondeceleratingthroughMach1.0.Thepitchtask duringdescentin followingtheVRI guidancewasfairly highin workload,andthehead-uppitchreference indicatorwasverysensitiveandindicatedfairlylargepitchresponsesfromverysmallpitchinputs.SincetheCG isbeingtransferredaftduringsupersonicdescent,frequentpitchtrimmingisrequired.A leveloff at9 km at Mach0.9wasaccomplishedwithoutdifficulty,andanITB (asdescribedabove)wascompleted.Furtherde- scentsasdirectedbyairtrafficcontrolplacedtheaircraftin thelandingpatternwith32tonsoffuel,6tonsabove theplannedamount. Fivetotalapproachesincludingthefinal full stoplandingwerecompleted.Theseincludeda straight-in localizeronlyapproachwiththecanardretracted;anoffsetapproachwiththenoseraiseduntilonfinal;amanual throttleoffsetapproach;amanualthrottlestraight-inapproach;andastraight-invisualapproachtoafull stop landing.Thefirst approachwiththecanardsretractedwasflownat360km/hrduetothelossof about12tonsof lift fromtheretractedcanards.Pitchcontrolwasnotasprecisein thisconfiguration.Therewasalsoalearning

35 curveeffectastheevaluationpilotgainedexperiencein makingverysmall,precisepitchinputswhichisneces- sarytoproperlyfly theaircraftonapproachandtoproperlyusethepitchreferenceindicator.Afterterminating theapproachat 60 m, a canardretracted,geardownlow passup therunwayat 30-40m wascompletedin accordancewithagroundeffectsexperimentrequirement.Thenose-upapproachdemonstratedthecapabilityto landthisaircraftwith thenoseretractedprovidingananglingapproachwith somesideslipis used.Theoffset approacheswerenotrepresentativeof thenormaloffsetapproachesflownin theHSRprogramsincetheyareto low approachonlyanddonottaxthepilot with thehighgainspotlandingtaskoutof thecorrectiveram.No untowardpitch/rollcouplingortendencytoovercontrolthepitchorroll axeswasnoted.Themanualapproaches wereveryinterestingin thattheTu-144LL,thoughaback-sidedairplaneon approach,wasnotdifficult to controlevenwiththehighlevelof throttlefrictionpresent.Theenginetimeconstantappearsreasonable.It was notedthatalargepitchingmomentresultsfrommoderateorgreaterthrottleinputswhichcanleadtoovercon- trollingthepitchaxisif thespeedisnottightlycontrolledandlargethrottleinputsarerequired.Thefull stop landingwasnotdifficultwithlightbrakingrequireddueto thedeceleratingeffectsof thedragparachutes.The flightterminatedwiththeevaluationpilottaxiingtheaircraftclearof therunwaytotheparkingarea.16tonsof fuelremained.

All testpointswereaccomplished,andseveraladditionaloptionaltestpointswerecompletedsincethe flight remainedaheadof theplannedfuel bum.Oneadditionalapproachwascompleted.Theplannedflight profilewasmatchedveryclosely,andallflight objectiveswereachieved.

36 Flight 23

Date of Flight: September 24, 1998 Flight Crew: Pilot in Command: Sergei Borisov Evaluation Pilot: Gordon Fullerton

Navigator: Victor Pedos Flight Engineer: Anatoli Kriulin Takeoff Time: 11:00 Local

Landing Time: 12:56 Local Flight Duration: 01:56 Takeoff Weight: 179 metric tons Landing Weight: 119 metric tons Landing Fuel: 16 metric tons Total Fuel Bum: 60 metric tons Takeoff CG: 40.8%

Landing CG: 40.6%

Flight Summary

Engines 2, 1, and 3 were started normally by the flight engineer. However, engine #4 temperature ap- proached a limit of 610 ° C so it was shut down. A slight tailwind condition existed. A successful start was made using cross-bleed air from the #3 engine. After engine start the aircraft was taxied for the brake wammp procedure and then into the lineup position. Power was set at 98 ° PLA, the start brake was released, and a nominal takeoff was made. The landing gear and canard were retracted on schedule, the nose was raised, and the aircraft was accelerated to an initial climb speed of 700 kin/hr. No special test points were planned during climb so that full attention could be devoted to flying the VRI profile as accurately as possible, to allow evaluation of the demanding pitch control task.

The aircraft was leveled at an altitude of 16.5 km and accelerated to Mach 2.0. A pitch capture maneuver of 2 ° nose up was flown, even though the many pitch adjustments required during the climb profile allowed a thorough examination of pitch control characteristics.

Next the highest priority test point of the supersonic cruise was accomplished: a set of frequency sweep maneuvers in the longitudinal, lateral, and directional axes.

After completion of the longitudinal frequency sweep at 700 km from the takeoff base the navigator called for a course reversal to the left. Lateral and directional sweeps were then completed. Control raps were accomplished in all axes, followed by steady heading sideslip maneuvers out to 4 ° of rudder deflection in each direction.

37 Justpriorto theplanneddescentgrossweightof 135tons,the#1throttlewasretardedto59°PLAandthe responseoftheaircraftnoted.Theasymmetrywastrimmedoutwithrudderandlateraltrimastheotherengines wereadvancedtomaximumthrust(115°PLA.)Aheadingcapturemaneuverwasflown.Speeddecreasedslowly toaboutMach1.9.

With32tonsof fuelremainingallenginesweresetto59° PLA,andtheaircraftwasallowedtodecelerate tointercepttheVRIprofilefor descent.Controlrapsin threeaxeswerecompletedpassingMach1.6.Leveloff wasatMach0.9and9kmaltitude.At thissubsoniccruiseconditiontheITBseriesof maneuversusedon the previousflightswascompleted. About200kmoutadescenttopattemaltitudewasbegunwith controlrapsmadepassingMach0.8Ap- proachingtheairfieldat500 km/hr the nose was lowered to 11°. About 15 km out and lined up with Runway 30, the nose was raised and the aircraft flown in a clean configuration over the runway at 100 m for a photo pass.

Turning left to downwind, the nose was lowered to 17 deg, the canard deployed, and the landing gear lowered. A visual approach was completed with manual control of thrust down to a go-around at 60 m.

On the downwind leg the #1 engine was retarded to 10 deg PLA, the landing gear lowered, and a three engine approach was flown, using the autothrottle system, with a three engine go-around initiated at 60 m.

The next pattern was set up with the canard retracted and using autothrottle, a descent was made leveling at 20 m above the runway. The autothrottle was disabled, and the aircraft kept level, maintaining 350 km/hr for about 10 sec for ground effects data.

The wind was reported at about 6 m/sec, above the limit of 2.5 m/sec, so the low-pass planned for Experi- ment 1.6 data was canceled.

The final pattern was begun with 16 tons fuel remaining. The standard configuration and procedure was used, with autothrottle engaged until about 5 m above the runway when the throttles were retarded to idle. After a smooth touchdown the nose gear was lowered, drag chutes deployed, and light braking brought the aircraft to taxi speed. The aircraft was parked and shutdown in the starmp area.

38 Form Approved REPORT DOCUMENTATION PAGE OMBNo.07704-0188

Public reporting burden for this collection of information is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources, gathering and maintaining the data needed, and completing and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of this collection of information, including suggestions for reducing this burden, to Washington Headquarters Services, Directorate for Information Operations and Reports, 1215 Jefferson Davis Highway, Suite 1204, Arlington, VA 22202-4302, and to the Office of Management and Budget, Paperwork Reduction Project (0704-0188), Washington, DC 20503.

1. AGENCY USE ONLY (Leave blank ]2. REPORT DATE 3. REPORTTYPE AND DATES COVERED I February 2000 Technical Memorandum 4. TITLE AND SUBTITLE 5. FUNDING NUMBERS A Qualitative Piloted Evaluation of the Tupolev Tu-144 Supersonic Transport 537-08-23

6. AUTHOR(S) Robert A. Rivers, E. Bruce Jackson, C. Gordon Fullerton, Timothy H. Cox, Norman H. Princen

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES) 8. PERFORMING ORGANIZATION REPORT NUMBER

NASA Langley Research Center L-17945 Hampton, VA 23681-2199

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES) 10. SPONSORING/MONITORING AGENCY REPORT NUMBER

National Aeronautics and Space Administration NASA/TM-2000-209850 Washington, DC 20546-0001

11. SUPPLEMENTARY NOTES Rivers and Jackson: Langley Research Center, Hampton, VA; Fullerton and Cox: Dryden Flight Research Center Edwards, CA; Princen: Boeing Commercial Airplane Group, Long Beach, CA

12a. DISTRIBUTION/AVAILABILITY STATEMENT 12b. DISTRIBUTION CODE

Unclassified-Unlimited Subject Category 08 Distribution: Standard Availability: NASA CASI (301) 621-0390

13. ABSTRACT (Maximum 200 words) Two U.S. research pilots evaluated the Tupolev Tu-144 supersonic transport aircraft on three dedicated flights: one subsonic and two supersonic profiles. The flight profiles and maneuvers were developed jointly by Tupolev and U.S. engineers. The vehicle was found to have unique operational and flight characteristics that serve as lessons for designers of future supersonic transport aircraft. Vehicle subsystems and observed characteristics are described as are flight test planning and ground monitoring facilities. Maneuver descriptions and extended pilot narratives for each flight are included as appendices.

14. SUBJECT TERMS 15. NUMBER OF PAGES HSR, flying qualities, supersonic transport, high-speed flight, Gromov, Tupolev, Zhuk- 44 ovsky, Tu- 144 16. PRICE CODE A03

17, SECURITY CLASSIFICATION 18, SECURITY CLASSIFICATION 19, SECURITY CLASSIFICATION 20, LIMITATION OF REPORT OF THIS PAGE OF ABSTRACT OF ABSTRACT Unclassified Unclassified Unclassified UL

NSN 7540-01-280-5500 Standard Form 298 (Rev. 2-89) Prescribed by ANSI Std. Z39-18 298-102